Vane assembly configured for turning a flow in a gas turbine engine, a stator component comprising the vane assembly, a gas turbine and an aircraft jet engine

ABSTRACT

A vane assembly configured for turning a flow in a gas turbine engine includes a stationary main guide vane and an additional guide vane, wherein a leading edge of the additional guide vane is positioned upstream of a leading edge of the main guide vane and wherein the additional guide vane extends a distance along the main guide vane towards a trailing edge of the main guide vane forming a passageway between the additional guide vane and the main guide vane.

BACKGROUND AND SUMMARY

The present invention relates to a vane assembly configured for turning a flow in a gas turbine engine. The invention is also related to a stator component comprising the vane assembly.

The gas turbine engine is especially intended for an aircraft jet engine. Jet engine is meant to include various types of engines, which admit air at relatively low velocity, heat it by combustion and shoot it out at a much higher velocity. Accommodated within the term jet engine are, for example, turbojet engines and turbo-fan engines.

In the aircraft jet engine, stationary guide vane assemblies are used to turn the flow from one angle to another. The stationary guide vane assembly may be applied in a stator component of a turbo-fan engine at a fan outlet, in a Turbine Exhaust Case (TEC) and even in an InterMediate Case (IMC).

The flow turning results in flow diffusion, i.e. pressure increase which puts serious limits on the amount of flow turning allowed for a given number of guide vanes. The number of guide vanes vs. the amount of flow turning often is governed by the so called solidity parameter “s”, i.e. the ratio between the vane chord “c” or length and the distance between two neighbouring vanes also called pitch “p”, “s=c/p”.

Turning the flow from 45 degrees to 0 degrees (axial direction) requires for instance a solidity of about 1.45. Depending on the flow Mach number and the radius of the component, this may result in more than 100 vanes. The design constraints are made even more challenging when thick structurally bearing vanes with engine servicing have to be used.

Typical Fan OGV configurations, which turn the flow by about 40-50 degrees, with structural loads but no servicing through may result in 58 vanes in the bypass duct. Typical TEC have about 14 vanes to turn the flow by about 30 degrees. TEC are however much thicker and have a max thickness to chord ratio of about 14%.

To achieve a larger amount of turning while keeping thick vanes, solutions have been presented based on high-lift devices which are comprised of a main vane and an additional vane just downstream to help the flow at turning further. These configurations are very much similar to high-lift wings on aircraft. However, the solutions proposed so far are often expensive and require a very long axial length to cope with the large thickness vs. flow turning. In aircraft engines, additional problems arise with thick and long vanes because of the upstream influence of the vane pressure field. Forcing of the upstream component becomes a major issue, especially in the engine core TEC and IMC. Other solutions opt for separation of mechanical and servicing functionality and aerodynamic functionality leading to the two rows solution with vanes followed by symmetrical struts.

It is desirable to achieve a vane assembly, which creates conditions for a large amount of flow turning while minimizing the upstream influence of the vane pressure field. Further, the axial extension of the vane assembly should be kept to a minimum.

According to an aspect of the present invention, a vane assembly is configured for turning a flow in a gas turbine engine comprising a stationary main guide vane and an additional guide vane, wherein a leading edge of the additional guide vane is positioned upstream of a leading edge of the main guide vane and wherein the additional guide vane extends a distance along the main guide vane towards a trailing edge of the main guide vane forming a passageway between the additional guide vane and the main guide vane. The main guide vane is preferably configured to turn an incoming flow and the additional guide vane is configured to assist the main guide vane in turning the incoming flow.

In other words, the additional stationary guide vane is arranged in the vicinity of the leading edge of the main guide vane and the additional guide vane is aerodynamically coupled to the main guide vane.

This design creates conditions for achieving a larger flow turning with a smaller number of main guide vanes (struts). For example, for a TEC, about 50% more flow turning may be achieved with about 30% less main guide vanes.

Further, a less complicated manufacturing (for example casting and forging) may be used compared to classical high lift devices.

Further, for IMC, fan OGV and TEC, this type of vane assembly may lead to more loading on upstream stages, a shorter engine length, reduced engine weight and reduced part count.

According to one embodiment of the invention, the additional guide vane is positioned relative to the main guide vane so that the passageway becomes more narrow in a downstream direction. Preferably, the additional guide vane is positioned relative to the main guide vane so that the passageway continuously narrows down from an upstream opening of the passageway to a downstream opening. More preferably, the additional guide vane is positioned relative to the main guide vane so that the passageway is shaped as a nozzle. Such a design creates conditions for tone noise reduction, the noise resulting from the vane interacting with upstream wakes, through phase cancellation effect and cavity volume dampening.

According to a further embodiment of the invention, the additional guide vane has an at least partly curved shape. Preferably, a first, upstream portion of a suction side of the additional guide vane extending from a leading edge of the additional guide vane is substantially straight and that a second, downstream portion of the suction side of the additional guide vane is curved. More preferably, the first, upstream portion of the suction side of the additional guide vane extends over at least 50% of the chord of the additional guide vane. Such a design creates conditions for reducing upstream pressure gradients.

According to a further embodiment of the invention, a leading edge of the additional guide vane has an elliptic cross sectional shape. In this way, upstream forcing may be reduced.

Further advantageous embodiments and advantages of the invention will be apparent from the following description, drawings and claims.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will be explained below, with reference to the embodiment shown on the appended drawings, wherein

FIG. 1 is a schematic side view of the engine cut along a plane in parallel with the rotational axis of the engine,

FIG. 2 is a perspective view of a stator component comprising an inventive guide vane assembly, and

FIG. 3 shows the vane assembly of FIG. 2 in cross section.

DETAILED DESCRIPTION

The invention will below be described for a high bypass ratio aircraft engine 100, see FIG. 1. The engine 100 comprises an outer housing 102, an inner housing 104 and an intermediate housing 106 which is concentric to the first two housings and divides the gap between them into an inner primary gas channel 108 for the compression of the propulsion gases and a secondary channel 110 in which the engine bypass circulates. Thus, each of the gas channels 108,110 is annular in a cross section perpendicular to an axial direction 112 of the engine 100.

The engine 100 comprises a fan 114 which receives ambient air 115, a booster or low pressure compressor (LPC) 116 and a high pressure compressor (HPC) 118 arranged in the primary gas channel 108, a combustor 120 which mixes fuel with the air pressurized by the high pressure compressor 118 for generating combustion gases which flow downstream through a high pressure turbine (HPT) 122 and a low pressure turbine (LPT) 124 from which the combustion gases are discharged from the engine.

A first or high pressure shaft joins the high pressure turbine 122 to the high pressure compressor 118 to substantially form a first or high pressure rotor. A second or low pressure shaft joins the low pressure turbine 124 to the low pressure compressor 116 to substantially form a second or low pressure rotor. The high pressure compressor 118, combustor 120 and high pressure turbine 122 are collectively referred to as a core engine. The second or low pressure shaft is at least in part rotatably disposed co-axially with and radially inwardly of the first or high pressure rotor.

The housings 102,104,106 are supported by structures 126 which connect the housings by radial arms. These arms are generally known as struts. The struts must be sufficiently resistant to provide this support and not to break or buckle in the event of a fan blade coming loose and colliding with them. Further, the struts are designed for transmission of loads in the engine.

Further, often the struts are hollow in order to house service components such as means for the intake and outtake of oil and/or air, for housing instruments, such as electrical and metallic cables for transfer of information concerning measured pressure and/or temperature, a drive shaft for a start engine etc. The struts can also be used to conduct a coolant.

The compressor structure 126 connecting the intermediate housing 106 and the inner housing 104 is conventionally referred to as an Intermediate Case (IMC) or Intermediate Compressor Case (ICC). The compressor structure 126 is designed for guiding the gas flow from the low pressure compressor 116 radially inwards toward to the high pressure compressor 118 inlet. The compressor structure 126 connecting the intermediate housing 106 and the inner housing 102 comprises a plurality of radial struts 208 see FIGS. 2 and 3, at mutual distances in the circumferential direction of the compressor structure. These struts 208 are structural parts, designed for transmission of both axial and radial loads and at least some are hollow in order to house service components.

FIG. 2 shows a perspective view of a stator component in the form of the compressor structure 126. The compressor structure 126 comprises an inner ring 202, an outer ring 204 encompassing the inner ring 202, and a plurality of vane assemblies 206 extending radially between the inner ring 202 and the outer ring 204. The vane assemblies 206 are circumferentially spaced and rigidly connected to the rings 202,204. Each of the vane assemblies 206 comprises the strut 208 and an additional guide vane 210, see also FIG. 3.

FIG. 3 shows one vane assembly 206 of FIG. 2 in an enlarged cross section view. The vane assembly 206 comprises a stationary main guide vane 208 (the strut) and the additional stationary guide vane 210. The main guide vane 208 is structurally bearing. The main guide vane 208 has a first sidewall 306 and a second sidewall 308 that are connected at a leading edge 310 and a trailing edge 312. The main guide vane 208 is configured to turn an incoming flow. The main guide vane 208 has an at least partly curved shape. The first side wall 306 of the main guide vane 208 is convex and defines a suction side. The second side wall 308 of the main guide vane 208 is concave and defines a pressure side.

The additional guide vane 210 has a first sidewall 314 and a second sidewall 316 that are connected at a leading edge 318 and a trailing edge 320. The additional guide vane 210 is configured to turn an incoming flow. The additional guide vane 210 has an at least partly curved shape. The first side wall 314 of the additional guide vane 210 is convex and defines a suction side. The second side wall 316 of the additional guide vane 210 is concave and defines a pressure side.

The magnitude of the turning of the gas flow in the stator component 126 depends on several parameters. In order to accomplish a turning of the gas flow in the magnitude of 40-60°, the main guide vane 208 has a cambered airfoil shape, see FIG. 3. In other words, the main guide vanes are designed with a sufficient curvature for a substantial turning of the gas flow. Hence, the main vane 208 is not only structural, but also has an aerodynamic function. More specifically, the direction of a mean camber line M at the leading edge 310 is inclined with an angle in relation to the direction of the mean camber line M at the trailing edge 312 corresponding to the desired turning angle. The direction of the mean camber line M at the leading edge 310 of the cambered main vane 208 is therefore inclined with at least 20°, suitably at least 30°, especially at least 40°, and preferably at least 50° in relation to the direction of the mean camber line M at the trailing edge 312.

The chord is defined as the distance between the leading edge 310 and the trailing edge 312 of the main vane 208 along the chord line C, see FIG. 3. The chord line C is defined as a straight line connecting the leading edge 310 and the trailing edge 312.

The thickness of the main vane 208 is defined as the maximum distance between the two opposing strut surfaces 306,308 in a direction perpendicular to a mean chamber line M. The mean camber line M is defined as the locus of points halfway between the upper and lower surfaces 306,308 of the main vane as measured perpendicular to the mean camber line itself. The camber A is defined as the maximum distance between the mean chamber line M and the chord line C measured perpendicular to the chord line. The main guide vane 208 has the shape of an airfoil in cross section. In other words, the mean camber line M is curved.

Further, the maximum thickness to chord ratio is another measure for the gas flow turning capacity of the struts. The maximum thickness is preferably less than 20%, especially less than 15% and more specifically about 10% of the chord according to the example shown in the drawings.

The leading edge 318 of the additional guide vane 210 is positioned upstream of the leading edge 310 of the main guide vane 208. The additional guide vane 210 extends a distance along the main guide vane 208 forming a passageway 322 between the additional guide vane 210 and the main guide vane 208. Thus, the additional guide vane 210 at least partly overlaps the main guide vane 208. The additional guide vane 210 is positioned relative to the main guide vane 208 so that the passageway 322 becomes more narrow in a downstream direction.

More specifically, the additional guide vane 210 is positioned relative to the main guide vane 208 so that the passageway 322 continuously narrows down from an upstream opening 324 of the passageway to a downstream opening 326. Thus, the leading edge 318 of the additional guide vane 210 is at a greater distance from a periphery of the main guide vane 208 than a trailing edge 320 of the additional guide vane 210 is from a periphery of the main guide vane 208. More specifically, the additional guide vane 210 is positioned relative to the main guide vane 208 so that the passageway 322 is shaped as a nozzle.

The trailing edge 320 of the additional guide vane 210 is arranged upstream of the trailing edge 312 of the main guide vane 208. More specifically, the trailing edge 326 of the additional guide vane 210 is arranged upstream of a point halfway of the chord of the main guide vane 208.

A pressure side of the additional guide vane 210 faces a suction side of the main guide vane 208, wherein the passageway 322 is defined therebetween.

The additional guide vane 210 has the shape of an airfoil in cross section. In other words, the mean camber line is curved.

Further, the additional guide vane 210 has a substantially smaller thickness to chord ratio than the main guide vane 208.

The vane assembly 206 is designed for turning a swirling gas flow. The swirling gas normally flows with an angle of 40-60° relative to the axial direction 112 of the engine. In this case the turning of the gas flow is in the combined axial-tangential and axial-radial directions.

A first, upstream portion 328 of a suction side 314 of the additional guide vane 210 extending from a leading edge 318 is substantially straight and a second, downstream portion 330 of the suction side 314 of the additional guide vane is curved. More specifically, the first, upstream portion 328 of the suction side of the additional guide vane extends over at least 50% of the chord of the additional guide vane. Especially, the first, upstream portion 328 of the suction side of the additional guide vane extends over about 70% of the chord of the additional guide vane.

A concave part of the additional guide vane 210 defines one side of the passageway 322. More specifically, the concave part of the additional guide vane 210 is arranged along a convex part of the main guide vane 208, wherein the passageway 322 is defined therebetween.

According to the shown embodiment, the vane assembly 206 is adapted to turn an incoming flow with a flow angle of about −45 degrees to an outgoing flow with a flow angle of about −5 degrees. It is expected that this type of vane assembly may turn a flow with about 60 degrees.

Further, it is expected that a minimum solidity of about 0.6 may be achieved with this type of vane assembly.

The struts are often hollow in order to house service components such as means for the intake and outtake of oil and/or air, for housing instruments, such as electrical and metallic cables for transfer of information concerning measured pressure and/or temperature etc. The struts normally have a symmetric airfoil shape in cross section in order to effect the gas flow as little as possible. The servicing requirement usually governs the number of struts required.

The invention is not in any way limited to the above described embodiments, instead a number of alternatives and modifications are possible without departing from the scope of the following claims.

The main guide vane 208, described above has an at least partly curved shape. Preferably, the mean camber line M is curved. However, according to an alternative, the main guide vane may be symmetrical in that its mean camber line may be straight. 

1. Vane assembly (206) configured for turning a flow in a gas turbine engine (100) comprising a stationary main guide vane (208) and an additional guide vane (210), wherein a leading edge (318) of the additional guide vane (210) is positioned upstream of a leading edge (310) of the main guide vane (208) and wherein the additional guide vane (210) extends a distance along the main guide vane (208) towards a trailing edge (312) of the main guide vane (208) forming a passageway (322) between the additional guide vane (210) and the main guide vane (208).
 2. Vane assembly according to claim 1, wherein the additional guide vane (210) is positioned relative to the main guide vane (208) so that the passageway (322) becomes more narrow in a downstream direction.
 3. Vane assembly according to claim 1 or 2, wherein the additional guide vane (210) is positioned relative to the main guide vane (208) so that the passageway (322) continuously narrows down from an upstream opening (324) of the passageway to a downstream opening (326).
 4. Vane assembly according to any preceding claim, wherein the additional guide vane (210) is positioned relative to the main guide vane (208) so that the passageway (322) is shaped as a nozzle.
 5. Vane assembly according to any preceding claim, wherein the additional guide vane (210) has an at least partly curved shape.
 6. Vane assembly according to claim 5, wherein a first, upstream portion of a suction side of the additional guide vane (210) extending from a leading edge of the additional guide vane (210) is substantially straight and that a second, downstream portion of the suction side of the additional guide vane (210) is curved.
 7. Vane assembly according to claim 6, wherein the first, upstream portion of the suction side of the additional guide vane (210) extends over at least 50% of the chord of the additional guide vane.
 8. Vane assembly according to claim 6, wherein the first, upstream portion of the suction side of the additional guide vane (210) extends over about 70% of the chord of the additional guide vane.
 9. Vane assembly according to any of claims 5-8, wherein a concave part of the additional guide vane (210) defines one side of the passageway.
 10. Vane assembly according to any preceding claim, wherein the main guide vane (208) has an at least partly curved shape.
 11. Vane assembly according to any preceding claim, wherein a concave part of the additional guide vane (210) is arranged along a convex part of the main guide vane (208), wherein the passageway is defined therebetween.
 12. Vane assembly according to any preceding claim, wherein a leading edge of the additional guide vane (210) has an elliptic cross sectional shape.
 13. Vane assembly according to any preceding claim, wherein the main guide vane (208) is configured to turn an incoming flow.
 14. Vane assembly according to any preceding claim, wherein the additional guide vane (210) is configured to turn an incoming flow.
 15. Vane assembly according to any preceding claim, wherein the main guide vane (208) has the shape of an airfoil in cross section.
 16. Vane assembly according to any preceding claim, wherein the additional guide vane (210) has the shape of an airfoil in cross section.
 17. Vane assembly according to claim 15 and 16, wherein a pressure side of the additional guide vane (210) faces a suction side of the main guide vane (208), wherein the passageway is defined therebetween.
 18. Vane assembly according to any preceding claim, wherein a trailing edge of the additional guide vane (210) is arranged upstream of the trailing edge of the main guide vane (208).
 19. Vane assembly according to any preceding claim, wherein a trailing edge of the additional guide vane (210) is arranged upstream of a point halfway of the chord of the main guide vane (208).
 20. Vane assembly according to any preceding claim, wherein the additional guide vane (210) has a substantially smaller thickness to chord ratio than the main guide vane (208).
 21. Vane assembly according to any preceding claim, wherein the main guide vane (208) is structurally bearing.
 22. Vane assembly according to any preceding claim, wherein the main guide vane (208) is adapted for housing service components.
 23. Vane assembly according to any preceding claim, wherein the additional guide vane (210) is stationary.
 24. Stator component comprising a plurality of said vane assemblies according to any preceding claim, wherein the main guide vanes (208) are circumferentially spaced.
 25. A gas turbine engine comprising a stator component according to claim
 24. 26. An aircraft jet engine comprising a stator component according to claim
 24. 